Integral sealing members for blades retained within a rotatable annular outer drum rotor in a turbomachine

ABSTRACT

A blade for a turbomachine includes a blade root portion for securing the blade to a rotatable annular outer drum rotor. The blade root portion includes one or more radial retention features for radially retaining each of the blade root portions within the rotatable annular outer drum rotor. Further, at least one of the radial retention feature(s) includes at least one sealing member integrated therewith.

GOVERNMENT SPONSORED RESEARCH

The project leading to this application has received funding from theClean Sky 2 Joint Undertaking under the European Union's Horizon 2020research and innovation program under grant agreement No. CS2-FRC-GAM2018/2019-807090.

FIELD

The present disclosure relates generally to turbomachines, and moreparticularly, to integral sealing members for blades retained within arotatable annular outer drum rotor of a turbomachine.

BACKGROUND

Gas turbine engines generally include a turbine section downstream of acombustion section that is rotatable with a compressor section to rotateand operate the gas turbine engine to generate power, such as propulsivethrust. General gas turbine engine design criteria often includeconflicting criteria that must be balanced or compromised, includingincreasing fuel efficiency, operational efficiency, and/or power outputwhile maintaining or reducing weight, part count, and/or packaging (i.e.axial and/or radial dimensions of the engine).

Gas turbine engines generally include a plurality of rotating rotorblades in at least one of a compressor of the compressor section or aturbine of the turbine section. Moreover, at least certain gas turbineengines also include a plurality of counter-rotating rotor blades in atleast one of the compressor of the compressor section or the turbine ofthe turbine section. Common rotating blades are assembled and retainedinternally by a disk, e.g. by means of the blade root, dovetail or firtree shaped or with a third part as a rivet or a bolted joint. The diskis typically located internally respect to the blade row. Bladesretained within the rotatable annular outer drum rotor are similarlyattached to a rotating part, but externally. As such, the outer drumrotor generally includes a drum, the blades, and separate rotatingshrouds.

For example, as shown in FIG. 1 , a cross-sectional view of acounter-rotating blade 1 according to conventional construction isillustrated. As shown, the blade 1 includes a blade root portion 2 forsecuring to a drum rotor 3. In addition, as shown, a separate rotatingshroud 4 may be attached to the drum rotor 3 for further securing theblade root portion within the drum rotor 3. In certain instances, aseparate sealing member 5 may be attached to the rotating shroud 4 toprevent undesirable air flow purge from entering the external cavitiesof the gas turbine engine, thereby limiting the loss of engineperformance. In addition, the sealing member 5 may act as a thermalbarrier, thereby providing thermal protection to the drum rotor 3.

Notwithstanding the aforementioned, there is a continuing need forimproved features associated with the blades retained within the drumrotor so as to improve operation efficiency of the gas turbine engine.Accordingly, the present disclosure is directed to an integrated sealingmember for blades retained within a routable annular outer drum rotor ofa turbomachine.

BRIEF DESCRIPTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In one aspect, the present disclosure is directed to a turbomachine. Theturbomachine includes a rotatable annular outer drum rotor connected toa first plurality of blades. Each of the first plurality of bladesincludes a blade root portion secured to the rotatable annular outerdrum rotor. Each of the blade root portions includes one or more radialretention features for radially retaining each of the blade rootportions within the rotatable annular outer drum rotor. Further, atleast one of the one or more radial retention features comprising atleast one sealing member integrated therewith.

In another aspect, the present disclosure is directed to a blade for aturbomachine. The blade includes a blade root portion for securing theblade to a rotatable annular outer drum rotor. The blade root portionincludes one or more radial retention features for radially retainingeach of the blade root portions within the rotatable annular outer drumrotor. Further, at least one of the one or more radial retentionfeatures includes at least one sealing member integrated therewith. Itshould be appreciated that the blade may further include any of theadditional features as described herein.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 illustrates a cross-sectional view of a turbine blade accordingto conventional construction;

FIG. 2 illustrates a longitudinal sectional view of one embodiment of anaircraft turbofan gas turbine engine with a counter rotating lowpressure turbine supported by an inter-turbine frame axially locatedbetween the low pressure turbine and a high pressure turbine;

FIG. 3 illustrates an enlarged view illustration of the inter-turbineframe and counter rotating low pressure turbine rotors of the engine inFIG. 1 ;

FIG. 4 illustrates an enlarged view illustration of a fan frame andforward bearings and sump of the engine in FIG. 1 ;

FIG. 5 illustrates a cross-sectional view of one embodiment of one ofthe first plurality of low pressure turbine blades according to thepresent disclosure;

FIG. 6 illustrates a perspective view of one embodiment of one of thefirst plurality of low pressure turbine blades according to the presentdisclosure;

FIG. 7 illustrates another perspective view of the low pressure turbineblade illustrated in FIG. 6 ;

FIG. 8 illustrates yet another perspective view of the low pressureturbine blade illustrated in FIG. 6 ;

FIG. 9A illustrates a simplified, front view of one embodiment of one ofthe plurality of low pressure turbine blades according to the presentdisclosure;

FIG. 9B illustrates a simplified, front view of another embodiment ofone of the plurality of low pressure turbine blades according to thepresent disclosure;

FIG. 9C illustrates a simplified, front view of yet another embodimentof one of the plurality of low pressure turbine blades according to thepresent disclosure;

FIG. 9D illustrates a simplified, front view of still another embodimentof one of the plurality of low pressure turbine blades according to thepresent disclosure;

FIG. 9E illustrates a simplified, front view of another embodiment ofone of the plurality of low pressure turbine blades according to thepresent disclosure;

FIG. 9F illustrates a simplified, front view of yet another embodimentof one of the plurality of low pressure turbine blades according to thepresent disclosure; and

FIG. 10 illustrates a top view of one embodiment of a sealing member fora low pressure turbine blade according to the present disclosure.

Repeat use of reference characters in the present specification anddrawings is intended to represent the same or analogous features orelements of the present invention.

DETAILED DESCRIPTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention.

As used herein, the terms “first”, “second”, and “third” may be usedinterchangeably to distinguish one component or feature from another andare not intended to signify location, importance, or magnitude of theindividual components or features.

The terms “forward” and “aft” refer to relative positions within a gasturbine engine or vehicle, and refer to the normal operational attitudeof the gas turbine engine or vehicle. For example, with regard to a gasturbine engine, forward refers to a position closer to an engine inletand aft refers to a position closer to an engine nozzle or exhaust. Theterms “upstream” and “downstream” refer to the relative direction withrespect to fluid flow in a fluid pathway. For example, “upstream” refersto the direction from which the fluid flows, and “downstream” refers tothe direction to which the fluid flows. The terms “coupled,” “fixed,”“attached to,” and the like refer to direct coupling, fixing, orattaching, as well as indirect coupling, fixing, or attaching throughone or more intermediate components or features, unless otherwisespecified herein. The singular forms “a”, “an”, and “the” include pluralreferences unless the context clearly dictates otherwise.

Approximating language, as used herein throughout the specification andclaims, is applied to modify any quantitative representation that couldpermissibly vary without resulting in a change in the basic function towhich it is related. Accordingly, a value modified by a term or terms,such as “about”, “approximately”, and “substantially”, are not to belimited to the precise value specified. In at least some instances, theapproximating language may correspond to the precision of an instrumentfor measuring the value, or the precision of the methods or machines forconstructing or manufacturing the components and/or systems. Forexample, the approximating language may refer to being within a 10percent margin.

Here and throughout the specification and claims, range limitations arecombined and interchanged, such ranges are identified and include allthe sub-ranges contained therein unless context or language indicatesotherwise. For example, all ranges disclosed herein are inclusive of theendpoints, and the endpoints are independently combinable with eachother.

Referring now to the drawings, FIG. 2 illustrates a schematic view ofone embodiment of turbomachine, such as a turbofan gas turbine engine 10according to the present disclosure. As shown, the gas turbine engine 10is circumscribed about an engine centerline 8 and includes a fan section12 which receives inlet airflow of ambient air 14. The fan section 12has counter rotating first and second fans 6 and 7 including first andsecond fan blade rows 13 and 15 and counter rotating first and secondboosters 16 and 17, respectively. The counter rotating first and secondboosters 16 and 17 are axially located between the counter rotatingfirst and second fan blade rows 13 and 15, an arrangement which providesreduced noise emanating from the fan section 12. Following the fansection 12 is a high pressure compressor (HPC) 18, a combustor 20 whichmixes fuel with the air 14 pressurized by the HPC 18 for generatingcombustion gases which flow downstream through a high pressure turbine(HPT) 24, and a counter rotating low pressure turbine (LPT) 26 fromwhich the combustion gases are discharged from the engine 10. The engine10 is designed such that the last stage of the second booster 17 and, inthe exemplary embodiment, the second fan blade row 15 are counterrotatable with respect to the high pressure compressor 18. This reducesthe sensitivity of the engine 10 to airflow inlet distortion of the fansection 12. It also reduces mutual sensitivity to rotating stall cellsin the other rotors.

A high pressure shaft 27 joins the HPT 24 to the HPC 18 to substantiallyform a first or high pressure rotor 33. The high pressure compressor 18,the combustor 20, and the high pressure turbine 24 collectively arereferred to as a core engine 25 which includes, for the purposes of thispatent, the high pressure shaft 27. The core engine 25 is modular suchthat as a single unit it can be independently replaced separate from theother parts of the gas turbine.

A bypass duct 21 radially, bounded by a fan casing 11 and a rotatableannular radially inner bypass duct wall 9, surrounds the counterrotating first and second boosters 16 and 17 and an inlet duct 19 to thehigh pressure compressor 18 of the core engine 25. The bypass duct 21 isradially bounded by a fan casing 11 and an annular radially inner bypassduct wall 9. The radially inner bypass duct wall 9 includes a rotatablewall section 22 fixedly mounted to the second fan blade row 15 and fromwhich the second booster 17 depends radially inwardly. A radially outerportion 23 of the second fan blade row is radially disposed within thebypass duct 21.

Referring to FIGS. 2 and 3 , the counter rotating low pressure turbine26 includes an annular outer drum rotor 136 rotatably mounted to a lowpressure inner shaft 130 by an aft low pressure inner conical shaftextension 132. The outer drum rotor 136 includes a plurality of lowpressure turbine blade rows 138 extending radially inwardly therefromand axially spaced from each other. The drum rotor 136 is cantileveredoff of a final stage 139 of the low pressure turbine blade rows 138which is bolted to the aft low pressure inner conical shaft extension132. The counter rotating low pressure turbine 26 also includes anannular low pressure inner drum rotor 146 rotatably mounted to a lowpressure outer shaft 140 by an aft low pressure outer conical shaftextension 142. The inner drum rotor 146 includes a plurality of secondlow pressure turbine blade rows 148 extending radially outwardlytherefrom and axially spaced from each other. The first low pressureturbine blade rows 138 are interdigitated with the second low pressureturbine blade rows 148.

The low pressure outer shaft 140 drivingly connects the inner drum rotor146 to the second fan blade row 15 and the second booster 17. The secondfan blade row 15 is connected to the low pressure outer shaft 140 by aforward conical outer shaft extension 143. The low pressure outer shaft140, the inner drum rotor 146, the second fan blade row 15, and thesecond booster 17 are major components of a low pressure outer rotor202. The low pressure inner shaft 130 drivingly connects the outer drumrotor 136 to the first fan blade row 13 and the first booster 16. Thefirst fan blade row 13 is connected to the low pressure inner shaft 130by a forward conical inner shaft extension 133. The low pressure innershaft 130, the outer drum rotor 136, the first fan blade row 13, and thefirst booster 16 are major components of a low pressure inner rotor 200.

The first booster 16 includes an annular first booster rotor section 166including the rotatable wall section 22 from which axially spaced apartfirst booster blade rows 168 extend radially inwardly. The annular firstbooster rotor section 166 is illustrated as being integrally bladed in amanner similar to an integrally bladed disk, commonly referred to as aBlisk, or an integrally bladed rotor which has been used in conventionalrotors because they are lightweight and allow no blade attachmentleakage. The operating low speeds of the boosters and the low weightintegrally bladed disk design of the first booster rotor section 166helps minimize stresses and deflections of the first booster rotorsection 166.

The second booster 17 includes an annular second booster rotor section170 from which axially spaced apart second booster blade rows 172 extendradially outwardly. A radially inner portion 28 of the second fan bladerow 15 is radially disposed within the inlet duct 19 and rotates withthe second booster 17 and therefore is considered part of the secondbooster 17 and a second booster blade row 172. The first and secondbooster blade rows 168 and 172 are interdigitated and are counterrotating. The first and second fan blade rows 13 and 15 are fixedlyattached to the first and second booster rotor sections 166 and 170,respectively. The low pressure inner and outer shafts 130 and 140,respectively, are at least, in part, rotatably disposed co-axially withand radially inwardly of the high pressure rotor 33.

The gas turbine engine 10 also has frame structure 32 including aforward or fan frame 34 connected by an engine casing 45 to a mid-engineor inter-turbine frame 60. The second fan blade row is axially locatedclose to struts 35 of the fan frame 34 and so the leading edges ofstruts 35 are swept or leaned axially aftwardly to reduce noise. Theengine 10 is mounted within or to an aircraft such as by a pylon (notillustrated) which extends downwardly from an aircraft wing. Theinter-turbine frame 60 includes a first structural ring 86, which may bea casing, disposed co-axially about the centerline 8. The inter-turbineframe 60 further includes a second structural ring 88 disposedco-axially with and radially spaced inwardly of the first structuralring 86 about the centerline 8. The second structural ring 88 may alsobe referred to as a hub. A plurality of circumferentially spaced apartstruts 90 extend radially between the first and second rings 86 and 88and are fixedly joined thereto. The struts 90 are hollow in theexemplary embodiment of the invention illustrated herein but, in otherembodiments, the struts may not be hollow. Because the inter-turbineframe 60 is axially located between the HPT 24 and the LPT 26 of thehigh pressure rotor 33 and the low pressure inner and outer rotors 200and 202, it is referred to as an inter-turbine frame also sometimesreferred to as a mid-engine frame. An inter-turbine transition duct 114between the HPT 24 and the LPT 26 passes through the inter-turbine frame60.

The engine is mounted to the aircraft at a forwardly located fan frameforward mount 118 on the fan frame 34 and at an aftwardly locatedturbine frame aft mount 120 on the inter-turbine frame 60. The engine 10may be mounted below an aircraft wing by a pylon at the forward mount118 and the aft mount 120 spaced axially downstream from the forwardmount 118. The aft mount 120 is used to fixedly join the inter-turbineframe 60 to a platform which is fixedly joined to the pylon. In oneembodiment, the aft mount 120 includes a U-shaped clevis 122.Conventional mounts often use a set of circumferentially spaced apartU-shaped clevises 122 (only one of the U-shaped clevises is shown in thecross-sectional illustrations in the figures) on the inter-turbine frame60. The U-shaped clevises 122 are designed to be connected by a set ofpins to a set of links. The links are connected to a platform on thebottom of the pylon. The U-shaped clevises 122 are one type of frameconnecting means for connecting the engine to an aircraft. Other typesof mounting means besides clevises are known in the aircraft industryand can be utilized to mount the frame of the present invention and theengine to the aircraft.

Referring more particularly to FIG. 4 , the low pressure outer rotor202, by way of the forward conical outer shaft extension 143, isrotatably supported axially and radially from the fan frame 34 by an aftthrust bearing 43 mounted in a first bearing support structure 44 and asecond bearing 36, a roller bearing, mounted in a second bearing supportstructure 47. The low pressure inner rotor 200, by way of the forwardconical inner shaft extension 133, is rotatably supported axially andradially from the fan frame 34 by a forward differential thrust bearing55 which is mounted between a forwardly extending extension 56 of theforward conical outer shaft extension 143 and the forward conical innershaft extension 133. The low pressure inner rotor 200 is furtherrotatably supported radially from the fan frame 34 by a forwarddifferential bearing 208, a roller bearing, between the low pressureinner shaft 130 and the low pressure outer shaft 140. The first andsecond bearing support structures 44 and 47 are fixedly attached to thefan frame 34.

Referring more particularly to FIG. 3 , the low pressure outer rotor202, by way of the aft low pressure outer conical shaft extension 142connected to the low pressure outer shaft 140, is rotatably supportedradially by a third bearing 76 within the inter-turbine frame 60. Thethird bearing 76 is disposed between an aft bearing support structure 97attached to an aft portion 110 of the inter-turbine frame 60 and aforward inner extension 190 of the aft low pressure outer conical shaftextension 142. The low pressure outer rotor 202 is most aftwardlyrotatably supported by the third bearing 76 which is thus referred to asan aftwardmost low pressure rotor support bearing. The inter-turbineframe 60 of the present invention is axially located between the HPT 24and the LPT 26 and thus substantially supports the entire low pressureturbine 26.

The low pressure inner rotor 200, by way of the aft low pressure innerconical shaft extension 132 connected to the low pressure inner shaft130, is rotatably supported radially by the aft low pressure outerconical shaft extension 142 of the low pressure outer rotor 202. Adifferential bearing 144 (also referred to as an inter-shaft bearing) isdisposed between an aft inner extension 192 of the aft low pressureouter conical shaft extension 142 and an outer extension 194 of the aftlow pressure inner conical shaft extension 132. This allows the lowpressure inner and outer rotors 200 and 202 to counter rotate.

Referring back to FIG. 2 , a forward high pressure end 70 of the highpressure compressor 18 of the high pressure rotor 33 is radiallyrotatably supported by a bearing assembly 80 mounted in a bearingassembly support structure 82 attached to the fan frame 34. Referringmore particularly to FIG. 2 , an aft end 92 of the high pressure rotor33 is aftwardly radially rotatably supported by a fifth bearing 94mounted in a forward bearing support structure 96 attached to a forwardportion 108 of the inter-turbine frame 60. The forward and aft bearingsupport structures 96 and 97 which are fixedly joined or attached to theforward and aft portions 108 and 110, respectively, of the inter-turbineframe 60 and thus are spaced axially apart. The forward and aft portions108 and 110, respectively, of the inter-turbine frame 60 are separatedby the second structural ring 88.

Forward and aft sump members 104 and 106 are joined to the inter-turbineframe 60 and carried by forward and aft bearing support structures 96and 97. The forward and aft sump members 104 and 106 support the fifthbearing 94 and the third bearing 76 in forward and aft cylindricalcentral bores 84 and 85, respectively, of the sump members. The fifthbearing 94 and the third bearing 76 have forward and aft fixed outerraces 176 and 178 that are fixedly connected to the forward and aftbearing support structures 96 and 97, respectively.

Located aft of the LPT 26 is an outlet guide vane assembly 150 whichsupports a stationary row of outlet guide vanes 152 that extend radiallyinwardly between a low pressure turbine casing 54 and an annular boxstructure 154. The outlet guide vane assembly 150 deswirls gas flowexiting the LPT 26. The low pressure turbine casing 54 connected isbolted to the engine casing 45 at the end of the inter-turbinetransition duct 114 between the HPT 24 and the LPT 26. A dome-shapedcover plate 156 is bolted to the annular box structure 154. The outletguide vane assembly 150 is not referred to and does not function as aframe because it does not rotatably support any of the engine's rotors.

The high pressure compressor 18 of turbofan gas turbine engine 10 of thepresent invention is operable and designed to operate with a relativelyhigh compressor pressure ratio in a range of about 15 to about 30 and anoverall pressure ratio in a range of about 40 to about 65. Thecompressor pressure ratio is a measure in the rise of pressure acrossjust the high pressure compressor 18. The overall pressure ratio is ameasure in the rise of pressure across the fan all the way through thehigh pressure compressor 18, i.e., it is a ratio of pressure exiting thehigh pressure compressor divided by pressure of ambient air 14 enteringthe fan section 12. The high pressure compressor 18 is illustratedhaving six high pressure stages 48 and three variable vane stages 50 forthe first four of the high pressure stages 48. Less than four variablevane stages 50 may be used. The high pressure compressor 18 has arelatively small number of the high pressure stages 48 and the inventioncontemplates using between 6 and 8 of the high pressure stages and aboutfour of the variable vane stages 50 or less. This makes for a shortengine while still having a high overall pressure ratio in a range of40-65.

The engine has a design bypass ratio in a range of 5-15 and a design fanpressure ratio in a range of 1.4-2.5. The counter rotating first andsecond fan blade rows 13 and 15 are designed to operate with tip speedsthat, for the two blade rows, sum to a range of about 1000 to 2500feet/sec which allows the use of light weight composite fan blades.Light weight, uncooled, high temperature capability, counter rotatingceramic matrix composite (CMC) airfoils may be used in the counterrotating low pressure turbine 26. Thus, the engine 10 and the fansection 12 may be described as having a sum of operational fan tipspeeds of the first and second fan blade rows 13 and 15 in a range of1000 to 2500 feet per second.

Referring still to FIG. 2 , a tip radius RT is illustrated, as measuredfrom the engine centerline 8 to a fan blade tip 188 of the first fanblade row 13 and a hub radius RH as measured from the engine centerline8 to a rotor hub 196 of the low pressure inner rotor 200 at an entrance186 to the inlet duct 19 to the high pressure compressor 18 of the coreengine 25. The engine 10 of the present invention may be designed with asmall fan inlet hub to tip radius ratio (RH/RT) in a range between 0.20and 0.35. For a given set of fan inlet and inlet duct annulus areas alow fan inlet hub to tip radius ratio allows a smaller fan diameter whencompared to a larger ratio.

However, fan inlet hub to tip radius ratio levels are constrained by theability to design a disk to support the rotating fan blades. The fanblades in the exemplary embodiment illustrated herein are made oflightweight composite materials or aluminum and rotor fan tip speeds aredesigned so that a fan disk 126 can be designed for the fan inlet hub totip radius ratio to be as low as 0.20. The low fan inlet hub to tipradius ratio allows low slopes and short lengths of the core enginetransition duct 124 between the fan section 12 and the high pressurecompressor 18 and of the inter-turbine transition duct 114 between theHPT 24 and the LPT 26.

Referring now to FIGS. 5-8 and 9A-9F, various views of multipleembodiments of one of the first plurality of low pressure turbine blades138 according to the present disclosure are illustrated. Referringparticularly to FIG. 5 , the low pressure turbine blade 138 includes ablade root portion 141 for securing to the annular outer drum rotor 136.In addition, each of the first plurality of low pressure turbine blades138 may include a blade tip portion 149 opposite the blade root portion141. Further, as shown, each of the blade root portions 141 may includeone or more structural radial retention features 145 for radiallyretaining each of the blade root portions 141 within the annular outerdrum rotor 136 and one or more axial retention features 147 for axiallyretaining each of the blade root portions 141 within the annular outerdrum rotor 136.

More specifically, as shown, the structural radial retention feature(s)145 may include a plurality of the radial retention hooks 155, 157 orflanges (e.g. mirrored hooks on opposing sides of the blade root portion141). For example, as shown generally in FIGS. 5-8 and 9A-9F, the radialretention feature(s) 145 may include opposing radial retention hooks155, 157. More specifically, as shown, the opposing radial retentionhooks 155, 157 may include, at least, a first radial retention hook 155and a second radial retention hook 157. In further embodiments, as shownparticularly in FIG. 9B, the radial retention hooks may further includea third radial retention hook 159, e.g. extending from the second radialretention hook 157 (or the first radial retention hook 155). In suchembodiments, the radial retention hooks 155, 157 are configured toprovide radial retention of the individual low pressure turbine blades138 (i.e. to prevent the blades 138 from falling out of the outer drumrotor 136). Further, as shown, outer surfaces of the radial retentionhooks 155, 157 are configured to provide radial retention of theindividual low pressure turbine blade 138 within the outer drum rotor136 during operation of the gas turbine engine 10.

In addition, as shown, at least one of the radial retention feature(s)145 includes at least one sealing member 160 integrated therewith. Insuch embodiments, as shown in FIGS. 5, 7, 8, and 9A-9F, the first radialretention hook 155 and/or the second radial retention hook 157 mayinclude the sealing member(s) 160 on an inner surface thereof. Morespecifically, as shown in FIGS. 9D and 9F, the sealing member(s) mayinclude a first sealing member 160 and a second sealing member 162. Insuch embodiments, as shown, the first and second radial retention hooks155, 157 may include the first and second sealing members 160, 162,respectively, on inner surfaces thereof. In further embodiments, asshown generally in FIGS. 5-8, 9A-9B and 9D-9F, the first radialretention hook 155 may define a first length, whereas the second radialretention hook 157 may define a second length. Moreover, as shown, thesecond length may be about one and half times as long as the firstlength or vice versa so as to provide a mounting location from thesealing member(s) 160, 162 described herein. In another embodiment, thesecond length may be about two or three times as long as the firstlength or vice versa.

In addition, as shown particularly in FIG. 5 and as mentionedpreviously, each of the blade root portions 141 may further include oneor more axial retention features 147 for axially retaining each of theblade root portions 141 within the rotatable annular outer drum rotor136. More specifically, as shown, the axial retention feature(s) 147 mayinclude a rotating shroud 158 attached to the rotatable annular outerdrum rotor 136. In such embodiments, the rotating shroud 158 and thesealing member(s) 160 may be separate features that are spaced apartfrom each other. For example, as shown generally in FIGS. 5-8 and 9A-9F,the sealing member(s) 160 is integrated with the turbine blade 138rather than being part of the rotating shroud 158. In addition, as shownin FIG. 10 , the sealing member(s) 160, 162 may have a honeycombconfiguration.

Referring particularly to FIGS. 9A-9F, the blade tip portions 149 forthe turbine blades 138 may include at least one additional sealingmember 164, 165. For example, as shown in FIGS. 9A-9D, the blade tipportions may have an arm member 174 with the additional sealing member164 secured thereto. In further embodiments, as shown in FIGS. 9E and9F, the blade tip portions 149 may have at least two arm members 174,175, each having at least one of the additional sealing members 164, 166secured thereto. In particular embodiments, as shown generally in FIGS.5-8 and 9A-9F, the sealing member(s) 160, 162 and the additional sealingmember(s) 164, 165 may include one or more steps.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

What is claimed is:
 1. A turbomachine defining an engine centerline,comprising: a rotatable annular outer drum rotor connected to a firstplurality of blades, with each of the first plurality of bladescomprising a blade root portion secured to the rotatable annular outerdrum rotor, each of the blade root portions comprising: one or moreradial retention features having a first radial retention hook forradially retaining each of the blade root portions within the rotatableannular outer drum rotor; and at least one sealing member provided alongand extending radially inwardly from, with respect to the enginecenterline, the one or more radial retention features, with the at leastone sealing member defined by a one or more steps; and at least oneshroud operably coupling the first radial retention hook to therotatable annular outer drum rotor and axially retaining the blade rootportions.
 2. The turbomachine of claim 1, wherein the turbomachinecomprises at least one of a turbine section, a compressor section, or agenerator.
 3. The turbomachine of claim 2, wherein the turbomachinecomprises the turbine section, the turbine section comprising a highpressure rotor comprising a high pressure turbine and a low pressureturbine comprising counter rotatable low pressure inner and outer rotorslocated aft of the high pressure rotor, the low pressure turbine furthercomprising the rotatable annular outer drum rotor connected to the firstplurality of blades and a rotatable annular inner drum rotor connectedto a second plurality of blades.
 4. The turbomachine of claim 1, whereinthe one or more radial retention features further comprises a firstradial retention feature and a second radial retention feature providedforward of the first radial retention feature.
 5. The turbomachine ofclaim 4, wherein the first radial retention feature includes the firstradial retention hook and the second radial retention feature includes asecond radial retention hook, and wherein at least one of the firstradial retention hook or the second radial retention hook comprises theat least one sealing member on an inner surface thereof.
 6. Theturbomachine of claim 5, further comprising a third radial retentionhook extending from the first radial retention feature.
 7. Theturbomachine of claim 5, wherein the at least one sealing member furthercomprises a first sealing member and a second sealing member, whereinthe first sealing member is provided on an inner surface of the firstradial retention hook, and the second sealing member is provided on aninner surface of the second radial retention hook.
 8. The turbomachineof claim 5, wherein the first radial retention hook defines a firstlength and the second radial retention hook defines a second length, thesecond length being at least one and a half times as long as the firstlength.
 9. The turbomachine of claim 1, wherein the at least one shroudis a rotating shroud of the rotatable annular outer drum rotor.
 10. Theturbomachine of claim 9, wherein the rotating shroud and the at leastone sealing member are spaced apart from each other.
 11. Theturbomachine of claim 1, wherein each of the first plurality of bladescomprises a blade tip portion opposite the blade root portion, the bladetip portion comprising at least one additional sealing member.
 12. Theturbomachine of claim 11, wherein the blade tip portion furthercomprises at least two arm members each comprising at least one of theadditional sealing members.
 13. The turbomachine of claim 11, whereinthe at least one additional sealing member comprise one or more steps.14. A blade for a turbomachine defining an engine centerline, the bladecomprising: a blade root portion for securing the blade to a rotatableannular outer drum rotor, and comprising: one or more radial retentionfeatures having a first radial retention hook for radially retaining theblade root portion within the rotatable annular outer drum rotor; and atleast one sealing member provided along and extending radially inwardlyfrom, with respect to the engine centerline, the one or more radialretention features, with the at least one sealing member defined by aone or more steps; wherein a shroud of the turbomachine operably couplesthe first radial retention hook to the rotatable annular outer drumrotor and axially retaining the blade root portion.
 15. The blade ofclaim 14, wherein the one or more radial retention features furthercomprise a first radial retention feature and a second radial retentionfeature provided forward of the first radial retention feature.
 16. Theblade of claim 15, wherein the first radial retention feature includesthe first radial retention hook and the second radial retention featureincludes a second radial retention hook, and wherein at least one of thefirst radial retention hook or the second radial retention hookcomprises the at least one sealing member on an inner surface thereof.17. The blade of claim 16, further comprising a third radial retentionhook extending from the second radial retention hook.
 18. The blade ofclaim 16, wherein the at least one sealing member further comprises afirst sealing member and a second sealing member, wherein the firstsealing member is provided on an inner surface of the first radialretention hook, and the second sealing member is provided on an innersurface of the second radial retention hook.
 19. The blade of claim 14,further comprising-a blade tip portion opposite the blade root portion,the blade tip portion comprising at least one additional sealing member.20. The turbomachine of claim 1, wherein the at least one shroud is aseparate component from the rotatable annular outer drum rotor and thefirst radial retention hook.